Wall for a turbomachine combustion chamber including an optimised arrangement of air inlet apertures

ABSTRACT

A rotationally symmetrical wall for an aircraft turbomachine combustion chamber, including primary holes, dilution holes positioned downstream from the primary holes, and a plug installation aperture upstream from the primary holes, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane (D) in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane (D), wherein the other primary holes are distributed equidistantly, at a predefined interval P, and the distance between the two primary holes is greater than the interval P separating two of the other adjacent primary holes.

TECHNICAL FIELD

The invention concerns a revolution wall for an aircraft turbomachine combustion chamber, including circumferential alignments of primary holes and of secondary holes.

STATE OF THE PRIOR ART

A turbomachine includes at least one turbine positioned at the outlet of a combustion chamber to extract energy from a primary flow of gases ejected by this combustion chamber, and to drive a compressor positioned upstream of the combustion chamber, and supplying this chamber with pressurised air.

The combustion chamber of a turbomachine includes two coaxial annular walls, which are respectively radially internal and radially external, and which extend from upstream to downstream, in the direction of the primary gas flow in the turbomachine, around the axis of the combustion chamber.

The annular walls are connected to one another at their upstream end by a back-of-chamber annular wall which extends roughly radially around the axis of the combustion chamber. This back-of-chamber annular wall is fitted with an annular row of injection systems which are regularly distributed around this axis of the combustion chamber to allow air and fuel to be introduced into the combustion chamber.

A combustion chamber is generally divided into an upstream internal region, commonly called the primary zone, and a downstream internal region, commonly called the dilution zone.

The primary zone is designed for the combustion of the air-fuel blend in roughly stoichiometric proportions. The dilution zone is designed to dilute and cool the gases resulting from the combustion in the primary zone, and to give the flow of these gases an optimum thermal profile for its transfer into the turbine positioned downstream of the combustion chamber.

To this end, each revolution wall of the combustion chamber includes first apertures, commonly called primary holes, arranged in a circumferential row, which are associated with the primary zone, and second air inlet apertures, commonly called dilution holes, which are arranged in a circumferential row, and which are associated with the dilution zone.

The combustion chamber also includes at least one ignition plug intended to initiate combustion of the air-fuel blend when the turbomachine is started.

The plug is installed through a plug installation aperture formed in the radially external annular wall of the chamber, in the axis of an injection system dedicated to ignition, or halfway, circumferentially, between two injection systems dedicated to ignition.

The plug installation aperture is positioned axially upstream and near the primary holes.

Each revolution wall also includes microperforations through which air is admitted into the chamber, in order to cool the wall.

Due to the small distance between the plug installation aperture and the primary holes, such microperforations cannot be made between the installation aperture and the adjacent primary holes, which reduces the effectiveness of the cooling of the wall in question.

ACCOUNT OF THE INVENTION

The invention proposes a revolution wall for an aircraft turbomachine combustion chamber, including a circumferential row of primary holes, a circumferential row of dilution holes positioned downstream from the said primary holes, and at least one ignition plug installation aperture positioned axially upstream from the primary holes, in which the primary holes, firstly, and the dilution holes, secondly, are distributed symmetrically either side of a median axial plane of the installation aperture, wherein each first primary hole, considered as one moves away circumferentially from the median axial plane in either direction, is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane,

characterized in that the said other primary holes are distributed equidistantly, at a predefined interval P, and in that the distance between the said first two primary holes is greater than the interval P separating the said other two adjacent primary holes.

Such an arrangement of the primary holes increases the space between the plug installation aperture and the adjacent primary holes. It is then possible to make holes known as microperforation holes between the plug installation aperture and the primary holes to increase cooling of the wall, without reducing the number of primary holes.

Preferably, each of the other primary holes is positioned at the same circumferential position as an associated dilution hole.

Preferably, the distance between the first primary hole and the second primary hole found as one moves away circumferentially from the median axial plane is less than the interval P separating two of the other said adjacent primary holes.

Preferably, the dilution holes are distributed equidistantly, at an interval roughly equal to half the said predefined interval P.

Preferably, the revolution wall includes multi-perforation holes located between the installation aperture and each of the said first primary holes.

An annular combustion chamber for an aircraft turbomachine, including a back-of-chamber annular wall and also two coaxial rotationally symmetrical walls, which are respectively internal and external, connected to the said back-of-chamber wall, and one of which at least is a revolution wall as previously defined.

The invention also concerns a turbomachine for aircraft, such as a turbojet or an aircraft turboprop engine, including an annular combustion chamber as previously defined.

BRIEF DESCRIPTION OF THE ILLUSTRATIONS

Other characteristics and advantages of the invention will come to light on reading the detailed description which follows, for the understanding of which reference will be made to the appended illustrations, in which:

FIG. 1 is a partial schematic axial section view of an aircraft turbomachine including a combustion chamber according to a first preferred embodiment of the invention;

FIG. 2 is a section schematic representation on a radial plane of a part of the wall of the combustion chamber showing the arrangement of the primary holes and of the plug installation aperture;

FIG. 3 is a developed representation of the portion of the wall of the combustion chamber represented in FIG. 2, showing the arrangement of the primary holes, the dilution holes and the plug installation aperture;

FIG. 4 is a detailed view on a larger scale of FIG. 3, showing the distribution of the multiperforations made in the wall.

DETAILED ACCOUNT OF PARTICULAR EMBODIMENTS

FIG. 1 represents an annular combustion chamber 10 of a turbomachine 12, such as an aircraft turbojet.

In a known manner, combustion chamber 10 is positioned downstream from a compressor (not represented) of turbomachine 12, which supplies combustion chamber 10 with a pressurised airflow 14. Combustion chamber 10 is also installed upstream from a turbine (not represented) of the turbomachine, intended to cause the compressor to rotate under the effect of the thrust of the gases coming from the combustion chamber.

Combustion chamber 10 includes two revolution walls coaxial with lengthways axis A of combustion chamber 10, a radially internal revolution wall 16 and a radially external revolution wall 18, which extend one inside the other.

These two revolution walls 16 and 18 are connected to one another at their upstream end by a back-of-chamber annular wall 26, in a known manner.

Back-of-chamber annular wall 26 includes an annular row of injection apertures 22 regularly distributed around axis A of combustion chamber 10, and in which injection systems 28 are installed.

Each injection system 28 is associated with a fuel injector 30, and is designed to deliver into combustion chamber 10 a portion of airflow 14 from the compressor and of the fuel from injectors 30.

Each of annular walls 16 and 18 includes two circumferential rows of air inlet apertures 32 and 34. These apertures 32, 34 are of globally circular section and are intended to inject another portion 36 of airflow 14 from the compressor into the combustion chamber.

In operation this other portion 36 of airflow 14 reaches air inlet apertures 32 and 34 through an annular bypass space 40 delimited by annular walls 16 and 18 of the combustion chamber, firstly, and casings 24 of turbomachine 12, surrounding combustion chamber 10, secondly.

A first row of apertures 32 is formed around an upstream region 42 of the combustion chamber, commonly called the primary zone, in which the combustion reactions of the air-fuel mix occur. The apertures 32 of this first row are for this reason commonly called primary holes.

A second row of apertures 34 is formed around a region 44 of combustion chamber 10, downstream from the primary zone, in which the combustion gases are diluted and cooled. This second region 44 is commonly called the dilution zone, and apertures 34 of this second row are for this reason commonly called dilution holes.

Turbomachine 12 also includes at least one ignition plug 46 emerging in primary zone 42 of combustion chamber 10 through a plug installation aperture 48 made in radially external wall 18 of combustion chamber 10.

Plug 46 is centred relative to a plane containing axis A of the combustion chamber and axis B of one of injection systems 28, where the said plane is the plane of FIG. 1.

Main axis C of ignition plug installation aperture 48 is slightly out-of-line towards an upstream direction relative to row of primary holes 32.

As can be seen in greater detail in FIGS. 2 to 4, primary holes 32 and dilution holes 34 are distributed either side of a median axial plane D relative to combustion chamber 10, passing through main axis C of installation aperture 48.

In the following description reference will be made to a half of each alignment 50, 52 of primary holes 32 and of dilution holes 34. It will be understood that the description of the other half of each alignment will be deduced by simple symmetry.

Alignment 50 of primary holes 32 includes, as one moves away circumferentially from median axial plane D, a first primary hole 321 and a second primary hole 322 which is further from median axial plane D than first primary hole 321.

In a similar manner, alignment 52 of dilution holes 34 includes, as one moves away circumferentially from median axial plane D, a first dilution hole 341, a second dilution hole 342 and a third dilution hole 343.

In this case the number of dilution holes 34 is twice the number of primary holes 32.

A dilution hole 340 is therefore centred on median axial plane D.

According to an unrepresented variant embodiment, the number of dilution holes 34 is equal to the number of primary holes 32.

Alignment 52 of dilution holes 34 then includes, as one moves away circumferentially from median axial plane D, a first dilution hole 341, a second dilution hole 342 and a third dilution hole 343; it does not include a dilution hole centred on median axial plane D.

Installation aperture 48 is positioned upstream from dilution holes 34, and upstream from primary holes 32 such that its main axis C is located close to primary holes 34.

The diameter of installation aperture 48 is thus relatively great, compared to the diameters of primary holes 32 and dilution holes 34.

To this end, according to the invention, first primary hole 321 is positioned at a circumferential position relative to median axial plane D which is between the circumferential position of first dilution hole 341 and of second dilution hole 342.

Each of the other primary holes 322, 32 is associated with a dilution hole 34 such that it is located at the same circumferential position as associated dilution hole 34.

Thus, according to the embodiment represented in the figures, for which the number of dilution holes 34 is equal to twice the number of primary holes 32, second primary hole 322 is located at the same circumferential position as third dilution hole 343.

According to the embodiment for which the number of dilution holes 34 is equal to the number of primary holes 32, where first primary hole 321 is located at a circumferential position relative to median axial plane D which is between the circumferential position of first dilution hole 341 and of second dilution hole 342. Second primary hole 322 is therefore located at the same circumferential position as second dilution hole 342.

According to another aspect of primary holes 32 and dilution holes 34, primary holes 32 other than first two primary holes 321 are distributed circumferentially equidistantly at a predefined interval P.

Dilution holes 34 which are associated with these other primary holes are thus distributed equidistantly at the same interval P.

In addition, according to the embodiment for which the number of dilution holes 34 is higher than the number of primary holes, the dilution holes are distributed at an interval equal to half the interval P separating two of the other primary holes 32.

Whatever the embodiment of the rotationally symmetrical wall, the first dilution hole is located at a circumferential distance from median axial plane D which is equal to half the interval P separating two of the other primary holes 32.

Second dilution hole 342 is located at a circumferential distance from median axial plane D which is equal to interval P separating two of the other primary holes 32 when the number of dilution holes is equal to twice the number of primary holes, or again at a circumferential distance from median axial plane D which is equal to one and a half times the value of interval P.

As was previously stated, first primary hole 321 is positioned circumferentially between first dilution hole 341 and second dilution hole 342.

Consequently, as can be seen in greater detail in FIG. 2, circumferential distance 56 between first primary hole 321 and median axial plane D is greater than half interval P, and less than one and a half times the value of interval P.

Circumferential distance 58 between first primary hole 321 and second primary hole 322 is less than the value of interval P.

Circumferential distance 60 between second primary hole 322 and the other adjacent primary hole 32 is equal to the value of interval P.

Circumferential distance 62 between first two primary holes 321, which are positioned symmetrically either side of median axial plane D, is thus equal to twice circumferential distance 56 between each first primary hole 321 and median axial plane D, and is greater than the value of interval P.

Circumferential distance 62 between the first two primary holes is such that they are relatively far from the plug installation aperture.

According to another aspect of revolution wall 18, as can be seen in FIG. 4, circumferential distance 56 between each first primary hole 321 and median axial plane D is sufficiently great that multiperforation holes 54 may be made in wall 18, in a zone located between installation aperture 48 and each first primary hole 321.

Multiperforation holes 54 enable wall 18 to be cooled whilst allowing air to flow from annular bypass space 40 towards the internal volume of combustion chamber 10.

These multiperforation holes 54 are distributed across the entire surface of wall 18, and not only in the zone located between installation aperture 48 and each first hole 321. 

1. A revolution wall for an aircraft turbomachine combustion chamber, comprising: a circumferential row of primary holes, a circumferential row of dilution holes arranged downstream from the primary holes, and at least one ignition plug installation aperture positioned axially upstream from the primary holes, wherein the primary holes and the dilution holes are distributed symmetrically either side of a median axial plane of the installation aperture, and wherein each first primary hole, considered as one moves away circumferentially from the median axial plane in either direction is positioned at a circumferential position between the circumferential positions of the first dilution hole and of the second dilution hole considered as one moves away circumferentially from the median axial plane, wherein other primary holes are distributed equidistantly, at an interval, and a distance between two primary holes is greater than the interval separating two adjacent primary holes.
 2. A revolution wall according to claim 1, wherein each of the other primary holes is positioned at the same circumferential position as an associated dilution hole.
 3. A revolution wall according to claim 1, wherein the distance between the first primary hole and the second primary hole found as one moves away circumferentially from the median axial plane is less than the interval separating two of said other adjacent primary holes.
 4. A revolution wall according to claim 1, wherein the dilution holes are distributed equidistantly, at an interval roughly equal to half said predefined interval.
 5. A revolution wall according to claim 1, further comprising multiperforation holes located between the ignition plug installation aperture and each of said first primary holes.
 6. An annular combustion chamber for an aircraft turbomachine, including a back-of-chamber annular wall and two coaxial revolution walls, which are respectively internal and external, connected to said back-of-chamber wall, and one of which at least is a revolution wall according to claim
 1. 7. A turbomachine for aircraft including an annular combustion chamber according to claim
 6. 